The present disclosure relates to a gas turbine engine and, more particularly, to a blade tip clearance control system therefor.
Gas turbine engines, such as those that power modern commercial and military aircraft, generally include a compressor to pressurize an airflow, a combustor for burning a hydrocarbon fuel in the presence of the pressurized air, and a turbine to extract energy from the resultant combustion gases.
The compressor and turbine sections include rotatable blade and stationary vane arrays. Within an engine case structure, the radial outermost tips of each blade array are positioned in close proximity to a shroud assembly. Outer air seals of the shroud assembly are located adjacent to the blade tips such that a radial tip clearance is defined therebetween.
During engine operation, the thermal environment in the engine varies and may cause thermal expansion or contraction. Such thermal expansion or contraction may not occur uniformly in magnitude or rate such that the radial tip clearance varies.
The radial tip clearance is typically designed so that the blade tips do not rub under high powered operations such as take-off when the blade disk and blades expand as a result of thermal expansion and centrifugal loads. When engine power is reduced to the cruise condition, the radial tip clearance increases.
To facilitate engine performance, at least some engines include a blade tip clearance control system to maintain a close radial tip clearance.